phase 1 for spacecraft design
TRANSCRIPT
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PHASE 1MANUFACTURING FEASIBILITY PHASE WITH PRIMARY
FOCUS ON OPTIMIZATION OF THE ULTRA-LIGHTWEIGHTCORE MATERIAL FABRICATION PROCESS AND PRODUCTION
OF SAMPLES FOR MECHANICAL PROPERTY ASSESSMENT BY
NASA.
BY
MAHESHWOR KC
SAI KUMAR BOMULE
SAI VEERENDRA POTLA
RAKESH BADRAJISYED ALI
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TARGET CORE PROPERTIES
Property Targeted core
material
properties
D en si ty , l b/ f t3
L es s t ha n 3
C el l S iz e, i n N /A
Compressive
Strength, psi
1000
Compressive
modulus, ksi
295
C r us h S t re n gt h , p s i 1 0 00
Plate shear
s tr en gt h ( L) , p si
640
Plate shear
m od ul us ( L) , k si
102
Plate shearstrength (W), psi
370
Plate shear
modulus (W), ksi
38
Co st of 4 X 8 X 1
panel
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INTRODUCTION
Increasing the core thickness
greatly increases the stiffness of
the honeycomb construction,
while the weight increase is
minimal.
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Flex core
Formability feasibility with reduced anticlastic
curvature and without buckling the cell walls.
Curvatures of very tight radii are easily formed.
Higher shear strength than hexagonal
honeycomb (when formed in tight radii).
HONEYCOMB CELL CONFIGURATION
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Aramid Fiber Honeycomb:
Use HRH-10 aramid fiber for core material.
dipped in a heat-resistant phenolic resin
Features:
High strength and toughness in a small cell size
Low density nonmetallic core.
CORE MATERIAL
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Table : Mechanical properties of HRH-10
MECHANICAL PROPERTIES OF HRH-10
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FABRICATION OF HONEYCOMB CORE
(STRENGTH OPTIMIZATION OF HRH-10)
Where,
Ef = the elasticity modulus of the skins
Ec = the in-plane elasticity modulus of the core
Tf = the thickness of the skins
Tc= the thickness of the core
B= the width of the beamd = is the distance between the centers of the two faces
e = the position of the neutral axis.
Figure : Typical cross section of
Honeycomb Sandwich
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Assumption: weak core and strong CFRP skin and using classic Bernoulli hypothesis
The position of the neutral axis
Flexural Rigidity
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MECHANICAL PROPERTIES AND TEST METHODS
1. Compressive strength test
Bare compressive test-without skin to core bonding
Stabilized compressive test- with skin to core bonding
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CRUSH STRENGTH
When the spacecraft using such sandwich panels is hit by hyper velocity
meteoroids the time while the panel going plastic deformation before
getting crushed will be crucial for the astronauts to get out of the space
crafts.
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2. Plate shear test method
The Shear strength and modulus values
dimension as 6" x 2" x 0.5" for non-metallic honeycomb fibers.
initiated by bonding the specimens to 1/2" thick steel loading plates.
loading rate is normally 0.020 inches per minute.
Shear deflections are to be measured with a displacement
transducer that senses the relative movement of the two plates.
Shear modulus is calculated from the slope of the initial
straight-line portion of the load-deflection curve.
Beam-flexure test can be used to test the core with higher
densities but our goal is used to design core with density
less than 3 pcf so we ignore this test.
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SELECTION OF SKIN
We made the decision to use IM10 12K epoxy composite with continuous
unidirectional fiber
Epoxy IM10 12K composite had high strength-to-weight and high stiffness-
to-weight properties and is stronger than steel, lighter than aluminum and
as stiff as titanium.
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Unidirectional prepreg tapes have been mostly used in the aerospace industry.
Fibers that have been oriented in 0/90 degree pattern to double the strength as
compared to unidirectional orientation.
use of only one layer of prepregs have higher chance of dimpling.
ORIENTATION OF FIBERS
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UNIDIRECTIONAL CFRP PREPREG TAPE
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The table above shows the tensile and compressive properties of IM10 12k at different orientations.
SKIN ORIENTATIONS MECHANICAL PROPERTIES
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REQUIREMENTS OF ADHESIVES
adhesives must function in adverse range of temperatures.
should resist the radiations and thus the micro cracking.
Insulation, high de-bonding energy.
Adhesive should pass the outgassing standard test
procedure according to ASTM E595.
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ASTM TEST FOR OUTGASSING
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SUMMARIZED RESULTS OF ASTM E595
If (Collected volatile condensable material)CVCM 0.10% and TML
(Total mass lost) 1.0%, the material passes.
If CVCM 0.10% and TML > 1.0%, the material can pass if TML WVR
1.0%.
If CVCM > 0.10%, the material fails.
If TML WVR (Water vapor regained) > 1.0%, the material fails.
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TWO PART EPOXIES
offer excellent cohesion, resist chemicals, bonds very well with most
materials
operate over adverse range of temperature (4K) to 550F.
meet or exceed NASAs outgassing specification.
~40 g/m2 adhesive per skin considered to optimize the strength-to-weight
ratio in bending of sandwich structure.
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ADHESIVE FILLETS
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CALCULATIONS
Proposed Thickness of the core (tc)= 0.70
Proposed Weight of the core(Wc): Density*Volume = 1.8 ft3*(0.7/12*10X11)= 11.55lb
For skin (IM 10, 12k)
Area covered by 2lb spool of IM 10=11*2800m=8420.8 ft2
Weight of 11X10 IM 10 for one face of the panel= 2*110/8420.8=0.026lbs
We propose to use 8 skins layer on each face of the core and 40g/m2 (0.008 lb/ ft2
adhesive per skin layer to get the optimal strength-to-weight ratio in bending of
sandwich structure.
Total weight of the two part epoxy used in the panel= 0.008*8 (gaps between
layers)*2(faces)*110(nominal area for the panel)=14.08 lbs
Total weight of the proposed honeycomb core sandwich panel= 0.026*16(total no.of layers)+14.08+11.55=26.046lbs
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CALCULATION CONTD
Density of the overall proposed sandwich panel= 26.046/9.167=2.84lb/ ft3
Overall density < 3lb/ ft3 (OK)
Flexural rigidity
From the table 2,
Diameter of the each tow (layer of skin)= Sqrt(0.18*4/3.14)=0.478mm=0.0188
Total thickness of skin on each side of the core (8 on each side)=8*0.0188=0.1504
Total depth of the proposed panel: (0.1504)*2+ 0.70= 1.0004~1
From Eq. 1
Assuming Ec
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CONCLUSIONS
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